This invention relates to a method and apparatus for detecting oscillatory faults in an aircraft flight control system. More particularly, this invention relates to a method and apparatus for detecting failures causing undesirable low frequency or high frequency (oscillatory) motion of an aerodynamic control-surface.
Commercial aircraft often include aerodynamic control surfaces. By altering the shape of such surfaces the flight response of the aircraft changes, enabling the pilot to alter the aircraft attitude. Typically, one or more sets of redundant actuators are coupled to a control surface to alter the surface displacement. The pilot or a primary flight computer issues commands which define actuator positions, and thus, the shape of an aerodynamic control-surface. Sensors monitor actuator positions and provide feedback signals to the primary flight computer enabling correction of actuator/surface position errors.
Faults occurring in an actuator control loop may result in undesirable oscillatory response of the actuator/surface. Typical faults leading to such oscillatory response include a worn actuator, high drift at an amplifier or a short-circuit in an electronic component. Other fault conditions also may cause oscillatory response. Such oscillation may fatigue the mechanical components (including the actuator), detract from aircraft ride quality and/or damage the aerodynamic control surface or other aircraft structure. Accordingly, there is a need to detect and correct such fault conditions.
As previously mentioned, an aircraft typically includes redundant actuators. Thus, in-flight correction involves shutting down the failed actuator and relying on the redundant actuator(s). Once on the ground the failed component is replaced.
A conventional oscillatory failure monitor is disclosed in U.S. Pat. No. 4,566,101 (Skonieczny et. al.). As disclosed therein, redundant sensors indicative of the same physical condition provide respective input signals to a monitor. When testing for fault conditions, the two input signals are compared to generate a difference signal which then is analyzed to determine if within a prescribed tolerance. For a transition of the difference signal from within-tolerance to out-of-tolerance an up-down counter is incremented. For a difference signal which stays within-tolerance for a prescribed time period, the up-down counter is decremented (not below a zero value). Thus, if an oscillatory fault occurs at one sensor causing the difference signal to go out-of-tolerance periodically, then the up-down counter increments (twice) for each oscillation period. To avoid registering transient errors as a fault condition, a fault is registered only when the up-down counter exceeds a prescribed count. One of the problems of such an approach is that the higher the oscillation frequency exhibited for a fault, the less the difference signal amplitude. This is because the amplitude for the oscillating (failed) sensor response typically decreases as frequency increases. Skonieczny et. al. describe a region 3 shown in their FIG. 3 in which faults are not detected even though beyond the relative threshold. Skonieczny et. al. recite that fault responses at such frequencies "may not be of concern". However, faults in such higher frequency region adversely impact aircraft ride quality and fatigue mechanical components. Accordingly, there is a need for a monitoring method and apparatus which detects both low frequency and high frequency fault conditions.